Nozzle segment for a gas turbine engine with ribs defining radially spaced internal cooling channels

Abstract

A nozzle segment for a gas turbine engine may generally include an airfoil having an exterior surface defining a pressure side and a suction side extending between leading and trailing edges. The airfoil may define an open internal volume within its interior for receiving a cooling medium. The open internal volume may include a primary internal cavity and a plurality of internal cooling channels in flow communication with the primary internal cavity. The primary internal cavity may extend within the interior of the airfoil from a location adjacent to the leading edge to a forward end of each of the internal cooling channels. The internal cooling channels may extend within the interior of the airfoil from the primary internal cavity towards the trailing edge. In addition, the internal cooling channels may be spaced apart radially by a plurality of ribs extending within the interior of the airfoil.

What is claimed is: 
     
       1. A nozzle segment for a gas turbine engine, the nozzle segment comprising:
 an airfoil formed from a ceramic matrix composite material, the airfoil including an exterior surface defining a pressure side and a suction side extending between a leading edge and a trailing edge, the airfoil defining an open internal volume within an interior of the airfoil for receiving a cooling medium, the open internal volume including a primary internal cavity and a plurality of internal cooling channels in flow communication with the primary internal cavity, the primary internal cavity extending within the interior of the airfoil from a location adjacent to the leading edge to a forward end of each of the plurality of internal cooling channels, the plurality of internal cooling channels extending within the interior of the airfoil from the primary internal cavity towards the trailing edge, wherein the plurality of internal cooling channels are spaced apart radially by a plurality of ribs extending within the interior of the airfoil; and 
 at least one channel member positioned within the interior of the airfoil, the at least one channel member defining an outer perimeter of each internal cooling channel, wherein the at least one channel member comprises a plurality of channel members, each channel member of the plurality of channel members defining one of the plurality of internal cooling channels, and wherein the plurality of channel members are positioned one on top of another within the interior of the airfoil, each rib being formed by adjacent radial ends of each pair of adjacent channel members of the plurality of channel members. 
 
     
     
       2. The nozzle segment of  claim 1 , wherein each rib is configured to extend in a chordwise direction between an adjacent pair of internal cooling channels of the plurality of internal cooling channels. 
     
     
       3. The nozzle segment of  claim 2 , wherein each internal cooling channel extends in the chordwise direction between the forward end and an aft end disposed between the forward end and the trailing edge, each rib being configured to extend in the chordwise direction between the adjacent pair of internal cooling channels such that the plurality of internal cooling channels are fluidly isolated from one another between the forward end and the aft end of each respective internal cooling channel of the plurality of internal cooling channels. 
     
     
       4. The nozzle segment of  claim 1 , wherein the at least one channel member is formed from one or more fabric plies. 
     
     
       5. The nozzle segment of  claim 1 , wherein the airfoil further defines at least one cooling hole extending between an internal cooling channel of the plurality of cooling channels and the exterior surface. 
     
     
       6. The nozzle segment of  claim 5 , wherein the at least one cooling hole is defined between the internal cooling channel and at least one of the pressure side or the suction side of the airfoil. 
     
     
       7. The nozzle segment of  claim 6 , wherein the at least one cooling hole comprises a first cooling hole and a second cooling hole, the first cooling hole being defined between the internal cooling channel and the pressure side of the airfoil, the second cooling hole being defined between the internal cooling channel and the suction side of the airfoil. 
     
     
       8. The nozzle segment of  claim 5 , wherein the at least one cooling hole is defined directly between the internal cooling channel and the trailing edge of the airfoil. 
     
     
       9. The nozzle segment of  claim 1 , wherein a radial thickness of each rib is less than a radial height of each of the plurality of internal cooling channels. 
     
     
       10. A nozzle segment for a gas turbine engine, the nozzle segment comprising:
 an airfoil formed from a ceramic matrix composite material, the airfoil including an exterior surface defining a pressure side and a suction side extending between a leading edge and a trailing edge, the airfoil defining an open internal volume within an interior of the airfoil for receiving a cooling medium, the open internal volume including a primary internal cavity extending within the interior of the airfoil in a chordwise direction from a location adjacent to the leading edge, the airfoil further including a plurality of ribs extending within the interior of the airfoil in the chordwise direction from the primary internal cavity towards the trailing edge, wherein the plurality of ribs are spaced apart from one another within the interior of the airfoil so as to define a plurality of radially spaced internal cooling channels in flow communication with the primary internal cavity; and 
 wherein each rib is formed from at least one channel member positioned within the interior of the airfoil, the at least one channel member defining an outer perimeter of each internal cooling channel, and wherein the at least one channel member comprises a plurality of channel members, each channel member of the plurality of channel members defining one of the plurality of internal cooling channels, and wherein the plurality of channel members are positioned one on top of another within the interior of the airfoil, each rib being formed by adjacent ends of each pair of adjacent channel members of the plurality of channel members. 
 
     
     
       11. The nozzle segment of  claim 10 , wherein each rib is configured to extend in the chordwise direction between an adjacent pair of internal cooling channels of the plurality of internal cooling channels. 
     
     
       12. The nozzle segment of  claim 10 , wherein the at least one channel member is formed from one or more fabric plies. 
     
     
       13. The nozzle segment of  claim 10 , wherein the airfoil further defines at least one cooling hole extending between an internal cooling channel of the plurality of cooling channels and the exterior surface. 
     
     
       14. A gas turbine engine, comprising:
 a compressor: 
 a combustor in flow communication with the compressor; and 
 a turbine in flow communication with the combustor, the turbine including a turbine nozzle formed from an annular array of nozzle segments, each nozzle segment comprising: 
 an airfoil formed from a ceramic matrix composite material, the airfoil including an exterior surface defining a pressure side and a suction side extending between a leading edge and a trailing edge, the airfoil defining an open internal volume within an interior of the airfoil for receiving a cooling medium, the open internal volume including a primary internal cavity and a plurality of internal cooling channels in flow communication with the primary internal cavity, the primary internal cavity extending within the interior of the airfoil from a location adjacent to the leading edge to a forward end of each of the plurality of internal cooling channels, the plurality of internal cooling channels extending within the interior of the airfoil from the primary internal cavity towards the trailing edge, 
 wherein the plurality of internal cooling channels are spaced apart radially by a plurality of ribs extending within the interior of the airfoil; and 
 wherein each rib is formed from at least one channel member positioned within the interior of the airfoil, the at least one channel member defining an outer perimeter of each internal cooling channel, and wherein the at least one channel member comprises a plurality of channel members, each channel member of the plurality of channel members defining one of the plurality of internal cooling channels, and wherein the plurality of channel members are positioned one on top of another within the interior of the airfoil, each rib being formed by adjacent ends of each pair of adjacent channel members of the plurality of channel members. 
 
     
     
       15. The gas turbine engine of  claim 14 , wherein each of the plurality of ribs define a radial passage that extends between adjacent internal cooling channels of the plurality of internal cooling channels.